With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
Compressed air from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
FIG. 2 shows an isometric view of a typical single stage cooled turbine in which there is a nozzle guide vane in flow series with a turbine rotor. The nozzle guide vane includes an aerofoil 31 which extends radially between inner 32 and outer 33 platforms. The turbine rotor includes a blade mounted to the peripheral edge of a rotating disc. The blade includes an aerofoil 32 which extends radially outwards from an inner platform. The radially outer end of the blade includes a shroud which sits within a seal segment 35. The seal segment is a stator component and attached to the engine casing. The arrows in FIG. 2 indicate cooling flows.
Internal convection and external films are the prime methods of cooling the gas path components—airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes (NGVs) consume the greatest amount of cooling air on high temperature engines. High-pressure blades typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.
The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K.
The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the extremities of the blades and vanes, and the working gas annulus endwalls.
In other examples, the turbine blades may be so-called shroudless blades in which there is no platform on the free end of the turbine blades. Such blades rotate radially inwards of a gas path wall commonly referred to as a seal segment. This is similar to the seal segment shown in FIG. 2 and includes a radially outer chamber which is provided with cooling air to keep the component cool during use. It is also well known to provide impingement cooling to the exterior of the gas path wall of a seal segment.
Cooling of the NGV end wall is achieved with the use of cooling air which is provided on the radial outer and radial inner of the gas path wall in appropriate chambers. From here the cooling air travels inside the vanes and through film cooling holes and the like as described above.
Typically, components of a gas turbine engine are metallic and cast and machined. Cavities may be cast during the casting of the piece, or machined in at a later date. These fabrication techniques generally mean that the geometry of the cavities need to be simple with the cooling air feed and exit holes created separately usually via secondary machining. Typical tolerances of these fabrications ultimately limit how small they can become and how closely they can mirror the base segment's shape.
Cast cavities can allow detailed features to be formed, however, because the ceramic cores used to make the cavities are prone to movement during the casting process they ultimately limit the smallest wall thickness that can be achieved which can result in unnecessarily thick walls and additional weight penalties.
Additionally, because the ceramic cores need to be held during the casting process the cavities either need to incorporate cores which project beyond and thus through the component wall, or are tied to other components with reasonably large ceramic bridges or vias. These bridges will interconnect the various cavities formed within the part in ways that may limit the ability to direct cooling flow between the various cavities in a controlled manner to enable efficient cooling function.
EP2369139 describes a nozzle segment for a gas turbine engine includes a flange which extends from a vane platform, the flange including a hollow cavity to reduce the weight of the component. The hollow cavity may include one or more purge openings.
The present invention seeks to provide an alternative air cooled component which can be fabricated by an additive layer manufacturing technique to provide an improved cooling functionality.